WO2015049446A1 - Chambre de combustion pour turbomachine a admission d'air homogene au travers de systemes d'injection de carburant - Google Patents
Chambre de combustion pour turbomachine a admission d'air homogene au travers de systemes d'injection de carburant Download PDFInfo
- Publication number
- WO2015049446A1 WO2015049446A1 PCT/FR2014/052446 FR2014052446W WO2015049446A1 WO 2015049446 A1 WO2015049446 A1 WO 2015049446A1 FR 2014052446 W FR2014052446 W FR 2014052446W WO 2015049446 A1 WO2015049446 A1 WO 2015049446A1
- Authority
- WO
- WIPO (PCT)
- Prior art keywords
- annular
- combustion chamber
- injector
- fairing
- air intake
- Prior art date
Links
- 238000002485 combustion reaction Methods 0.000 title claims abstract description 82
- 238000002347 injection Methods 0.000 title claims abstract description 69
- 239000007924 injection Substances 0.000 title claims abstract description 69
- 239000000446 fuel Substances 0.000 claims abstract description 35
- 238000011144 upstream manufacturing Methods 0.000 claims abstract description 21
- 230000002093 peripheral effect Effects 0.000 claims description 7
- 239000007789 gas Substances 0.000 description 4
- 230000000694 effects Effects 0.000 description 1
- 239000003344 environmental pollutant Substances 0.000 description 1
- 231100000719 pollutant Toxicity 0.000 description 1
- 239000000243 solution Substances 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/10—Air inlet arrangements for primary air
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/22—Fuel supply systems
- F02C7/222—Fuel flow conduits, e.g. manifolds
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/002—Wall structures
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/06—Arrangement of apertures along the flame tube
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/10—Air inlet arrangements for primary air
- F23R3/12—Air inlet arrangements for primary air inducing a vortex
- F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/283—Attaching or cooling of fuel injecting means including supports for fuel injectors, stems, or lances
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/35—Combustors or associated equipment
Definitions
- the present invention relates to an annular combustion chamber for a turbomachine, in particular for an aircraft propulsion unit.
- the invention applies in particular but not exclusively to combustion chambers provided with an annular row of fuel injectors, each comprising a head provided with a central fuel injection nose and a peripheral device for fuel injection.
- fuel injection for example multipoint type.
- injectors of this type are used in so-called "staged combustion" combustion chambers.
- the invention also relates to a combustion chamber module and a turbomachine comprising such a combustion chamber module.
- FIG. 1 illustrates a typical example of a turbomachine 1 of a known type, for example a double-body, dual-flow aircraft turbojet engine.
- the turbomachine 1 comprises, successively in the thrust direction represented by the arrow 2 also corresponding to the general flow direction of the gases in the turbomachine, a low-pressure compressor 4, a high-pressure compressor 6, an annular combustion chamber 8 , a high pressure turbine 10 and a low pressure turbine 11.
- the upstream and downstream directions are defined relative to the general direction of gas flow within the combustion chamber and more generally of the turbomachine.
- the combustion chamber 8 is mounted downstream of the high-pressure compressor 6 intended to supply this chamber with pressurized air, and upstream of the high-pressure turbine 10 intended to drive the high-pressure compressor in rotation. under the effect of the thrust of the gases from the combustion chamber.
- Figure 2 illustrates on a larger scale the combustion chamber 8 and its immediate environment.
- the combustion chamber 8 comprises two coaxial annular walls, respectively radially inner 12 and radially outer 13, which extend around the longitudinal axis 14 of the combustion chamber.
- annular walls 12 and 13 are attached downstream to inner casings 15 and outer 16 of the chamber, and are connected to each other at their upstream end by an annular bottom wall 18 of the combustion chamber.
- the annular bottom wall 18 comprises an annular row of orifices regularly distributed around the axis 14 of the combustion chamber, and in which are mounted injection systems 20 in which respective heads 21 of fuel injectors 22 are adjusted respectively. These fuel injectors 22 each have a fuel emission axis which merges with an axis 24 of the corresponding injection system 20.
- the injection systems 20 are mounted in the bottom wall 18 so as to be able to move slightly in a direction orthogonal to the axis 24 and thus tolerate the differential expansions affecting the combustion chamber 8, the injection systems 20 and casings 15 and 16, in operation.
- the assembly formed by the combustion chamber 8 and the fuel injectors 22 is called "combustion chamber module" in the present description.
- Each injection system 20 has an upstream end forming a bushing 26, a downstream end taking the form of a flared bowl 28 opening into the combustion chamber 8, and an annular air inlet 30 arranged between the bushing 26 and the bowl 28 and intended for the admission of a portion 31 of the air flow 32 from a diffuser 34 mounted at the outlet of the high pressure compressor of the turbomachine, with a view to a premixing of the intake air with the fuel from the fuel injector 22 mounted in the sleeve 26, within the injection system.
- the annular air inlet 30 is traversed by fins 36 for printing to the flow of air through them a gyration movement.
- the air inlet is thus of the type commonly known as "spin".
- annular walls 12 and 13 of the combustion chamber are connected at their upstream end to an annular fairing 40 having orifices 42 arranged facing the injection systems 20 for the passage of the fuel injectors 22 and the
- the main function of the fairing 40 is to protect the bottom wall 18 of the combustion chamber and to guide portions 44 and 46 of the airstream 32 flowing downstream. respectively along the inner annular walls 12 and outer 13 of the combustion chamber, within two internally 48 and outer 50 bypass spaces respectively.
- these parts 44 and 46 of the air flow 32 are respectively called "Internal bypass airflow" and "external bypass airflow”.
- the internal bypass 48 and outer 50 spaces form, with an upstream space 52 which connects them to one another, an enclosure in which the combustion chamber extends.
- each orifice 42 is located in upstream of the annular air inlet 30 relative to the axis 24 of the corresponding injection system.
- the air supply of the annular air inlet 30 of the injection systems 20 has an inhomogeneous nature around the axis 24 of each injection system, such as to induce a reduction in the performance of the chamber combustion in particular in terms of limitation of pollutant emissions and in terms of control of the thermal profile of the exhaust gas leaving the combustion chamber.
- each injector 22 has a fuel injection central nose 54, an axial air intake device 56 arranged around the central nose 54, and a peripheral fuel injection device 58 arranged around the fuel injection device.
- the fuel from the central nose 54 mixes, within each injection system, with the air admitted through the axial air intake device 56, while the fuel from the peripheral device In each injection system, fuel injection 58 is mixed with the air admitted through the annular air inlet 30 of the injection system.
- the injection systems of this type have a relatively large transverse size, likely to increase the inhomogeneities of the air flow 31.
- these injection systems require a relatively large air flow to operate, this which tends to accentuate the heterogeneities in the combustion chamber 8.
- a radially outer portion of the annular air inlet 30 of each injection system 20 receives a direct air flow 31a while a radially inner portion of the annular inlet 30 receives an indirect airflow 31b.
- the invention aims in particular to provide a simple, economical and effective solution to these problems, to avoid at least partly the aforementioned drawbacks.
- the invention proposes for this purpose an annular combustion chamber for a turbomachine, comprising:
- annular bottom wall provided with a plurality of injection systems each centered on a respective axis and each having an upstream end forming a socket intended to receive a head of a fuel injector, a downstream end opening into said chamber; combustion, and an annular air inlet arranged between said upstream and downstream ends so that the air admitted through said annular air inlet mixes, within the injection system, with the fuel from the 'fuel injector, and
- annular fairing covering an upstream side of said bottom wall and comprising a plurality of injector passage orifices respectively arranged facing said injection systems, said annular fairing and said bottom wall jointly delimiting an annular space in which opens the annular air inlet of each injection system.
- said annular fairing comprises a plurality of air intake orifices separated from said injector orifices.
- said sleeve of each of said injection systems passes through the corresponding injector orifice of said annular fairing and comprises at its upstream end an annular flange having a free end remote from said axis of the injection system by a first distance greater than or equal to a second distance separating an edge of said corresponding injector orifice and said axis of the injection system.
- annular flange of the bushing of each injection system makes it possible to mask the inlet of the corresponding injector orifice of the annular fairing and thus to substantially reduce to zero the flow of air supplying the air inlet.
- annular injection system via said injector passage orifice The supply of the annular air inlet is thus almost exclusively provided indirectly by air passing through the air intake orifices of the annular fairing.
- this configuration makes it possible to maintain the mobile character of each injection system with respect to the annular fairing and the bottom wall of the combustion chamber.
- said air intake orifices and said injector passage orifices are distributed so that at least one air intake port is arranged circumferentially between each pair of consecutive injector orifices along the circumference of said annular fairing.
- said air intake orifices are preferably alternately distributed with said injector passage orifices along the circumference of said annular fairing.
- the invention also relates to a combustion chamber module for a turbomachine, comprising:
- annular combustion chamber of the type described above, and an annular row of fuel injectors comprising respective injector heads mounted fitted respectively in said bushings of the injection systems of said combustion chamber.
- each injector head comprises a central fuel injection nose, an axial air intake device arranged around said central nose, and a peripheral fuel injection device arranged around said axial air intake device.
- said injector passage orifices of said annular fairing advantageously have respective isobarecentres which are inscribed on a first circle centered on an axis of said combustion chamber and having a first diameter.
- said air intake orifices of said annular fairing have respective isobarecentres which are inscribed on a second circle centered on the axis of said combustion chamber and having a second diameter. strictly greater than said first diameter of said first circle.
- said air intake orifices of said annular fairing have respective isobarycentres which are inscribed on said first circle.
- the invention finally relates to a turbomachine for an aircraft, comprising a combustion chamber module of the type described above.
- FIG. 1, already described is a partial schematic view in axial section of a turbomachine of a known type
- FIG. 2 already described, is a partial schematic view in axial section of a combustion chamber module of the turbomachine of FIG. 1, comprising an axial diffuser;
- FIG. 3 is a partial schematic view in axial section of a combustion chamber module of a turbomachine of a known type, comprising a centrifugal diffuser;
- FIG. 4 is a partial schematic view in axial section of a combustion chamber module of a turbomachine according to a first preferred embodiment of the invention
- FIG. 5 is a partial schematic view in axial section of a combustion chamber belonging to the combustion chamber module of Figure 4 and shown isolated;
- FIG. 6 is a partial schematic view of the combustion chamber module of FIG. 4, seen from upstream;
- FIG. 7 is a view similar to FIG. 6, illustrating an alternative embodiment of the combustion chamber module of FIG. 4;
- FIG. 8 is a view similar to FIG. 6, illustrating a combustion chamber module of a turbomachine according to a second preferred embodiment of the invention.
- FIGS. 4 to 6 illustrate a portion of a combustion chamber module 59 according to a first preferred embodiment of the invention.
- This combustion chamber module is part of a turbomachine whose Other parts may be of a conventional type, as illustrated in Figure 1 described above.
- Figures 4 and 6 show more particularly a rear portion of the combustion chamber 8 and the injectors 22 of the combustion chamber module, while Figure 5 illustrates the rear portion of the combustion chamber 8 alone.
- the annular fairing 40 ' which covers the upstream side of the combustion chamber 8 has a plurality of air intake orifices 60 separated from the injector orifices 42.
- the air intake ports 60 are alternately distributed with the injector passage holes 42 along the circumference of the annular fairing 40 '.
- each orifice for injector passage 42 is located upstream of the annular air inlet 30 with respect to the axis 24 of the corresponding injection system.
- each injection system 20 passes through the corresponding nozzle orifice 42 of the annular shroud 40'.
- the sleeve 26 comprises at its upstream end an annular flange 62.
- This annular collar 62 has a free end 64 remote from the axis 24 of the injection system 20 by a first distance d1 (FIG. 5) greater than or equal to a second distance d2 separating an edge of said injector passage orifice 42 and the axis 24 of the injection system.
- the annular flange 62 does not have symmetry of revolution. Indeed, the first distance d1 varies slightly around the axis 24 of the injection system.
- a radially outer portion of the annular flange 62 is wider than a radially inner portion thereof.
- the radially side The outer end 66 of the free end 64 is further apart from the axis 24 of the injection system than the radially inner side 66 'of the free end 64.
- the injector orifice 42 has no symmetry of revolution, so that the second distance d2 varies slightly about the axis 24 of the injection system.
- the above inequality between the first distance d1 and the second distance d2 is valid within each axial section plane of the combustion chamber module.
- each injector head has a fuel injection central nose 54, an axial air intake device 56 arranged around said central nose 54, and a peripheral fuel injection device 58 arranged around said fuel injection device 56.
- axial air intake is for example of the "multipoint" type, that is to say having an annular row of fuel ejection orifice.
- the injector passage orifices 42 of the annular fairing 40 have respective isobarecentres 68 which register on a first circle 70 centered on the axis 14 of the combustion chamber 8 and having a first diameter Dl.
- the air intake orifices 60 of the annular shroud 40 have respective isobarecentres 72 which are inscribed on a second circle 74 centered on the axis 14 of the combustion chamber. 8 and having a second diameter D2 strictly greater than the first diameter D1 of said first circle 70.
- the air intake ports 60 are offset radially outwardly of the annular fairing 40 '. This configuration is particularly advantageous when the diffuser supplying air to the combustion chamber is of the centrifugal type, as in the example of the prior art illustrated in FIG.
- the air intake ports 60 have an oblong shape in the circumferential direction.
- each air intake port 60 is remote from the first circle 70 above.
- each air inlet 60 may extend to the first circle 70, as shown in FIG. 7.
- the edge of each air intake port 60 advantageously has curved lateral regions so as to substantially follow the curve of the injector orifice 42 located nearby.
- the air intake orifices 60 of the annular fairing 40 have respective isobarecentres 72 which are inscribed on the first circle 70.
- This configuration is particularly advantageous when the diffuser supplying air to the combustion chamber is of the axial type, as in the example of the prior art illustrated in FIG.
- the air supplying the annular air inlet 30 of each injection system 20 passes exclusively or almost exclusively through the air intake orifices 60 of the annular fairing 40 '.
- the annular flange 62 of the bushing of each injection system 20 substantially prevents the passage of air around each injection system through the corresponding nozzle orifice 42.
- the annular flange 62 and the edge of the injector passage orifice 42 form an annular baffle for the air flow from the diffuser supplying the combustion chamber with pressurized air.
- the air supplying the annular air inlet 30 of each injection system 20 circulates beforehand by swirling within a space 78 (FIG. 4) defined between the bottom wall 18 and the annular fairing 40. 'of the combustion chamber. This results in an improvement in the homogeneity of the air supply of the annular air inlet 30 around its respective axis.
- the injector and air intake orifices are distributed alternately.
- the air intake orifices 60 and the injector orifices 42 are preferably distributed so that at least one air inlet 60 is circumferentially arranged between each pair of consecutive injector ports 42 along the circumference of the annular fairing 40 '.
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Fuel-Injection Apparatus (AREA)
- Pressure-Spray And Ultrasonic-Wave- Spray Burners (AREA)
- Combustion Methods Of Internal-Combustion Engines (AREA)
Abstract
Description
Claims
Priority Applications (6)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP14787226.1A EP3039342B1 (fr) | 2013-10-01 | 2014-09-29 | Chambre de combustion pour turbomachine a admission d'air homogène au travers de systèmes d'injection de carburant |
US15/023,954 US10180256B2 (en) | 2013-10-01 | 2014-09-29 | Combustion chamber for a turbine engine with homogeneous air intake through fuel injection system |
CN201480054491.6A CN105593602B (zh) | 2013-10-01 | 2014-09-29 | 通过燃料喷射系统均匀进气的涡轮发动机的燃烧室 |
RU2016116529A RU2660729C2 (ru) | 2013-10-01 | 2014-09-29 | Камера сгорания для турбинного двигателя с равномерным забором воздуха через систему впрыска топлива |
BR112016006688-0A BR112016006688B1 (pt) | 2013-10-01 | 2014-09-29 | Câmara de combustão anular para turbomáquina com admissão de ar homogêneo através de sistemas de injeção de carburante, módulo de câmara de combustão e turbomáquina para aeronave |
CA2925441A CA2925441C (fr) | 2013-10-01 | 2014-09-29 | Chambre de combustion pour turbomachine a admission d'air homogene au travers de systemes d'injection de carburant |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
FR1359503 | 2013-10-01 | ||
FR1359503A FR3011317B1 (fr) | 2013-10-01 | 2013-10-01 | Chambre de combustion pour turbomachine a admission d'air homogene au travers de systemes d'injection |
Publications (1)
Publication Number | Publication Date |
---|---|
WO2015049446A1 true WO2015049446A1 (fr) | 2015-04-09 |
Family
ID=49551693
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/FR2014/052446 WO2015049446A1 (fr) | 2013-10-01 | 2014-09-29 | Chambre de combustion pour turbomachine a admission d'air homogene au travers de systemes d'injection de carburant |
Country Status (8)
Country | Link |
---|---|
US (1) | US10180256B2 (fr) |
EP (1) | EP3039342B1 (fr) |
CN (1) | CN105593602B (fr) |
BR (1) | BR112016006688B1 (fr) |
CA (1) | CA2925441C (fr) |
FR (1) | FR3011317B1 (fr) |
RU (1) | RU2660729C2 (fr) |
WO (1) | WO2015049446A1 (fr) |
Families Citing this family (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JP5924618B2 (ja) * | 2012-06-07 | 2016-05-25 | 川崎重工業株式会社 | 燃料噴射装置 |
GB2548585B (en) * | 2016-03-22 | 2020-05-27 | Rolls Royce Plc | A combustion chamber assembly |
FR3070198B1 (fr) * | 2017-08-21 | 2019-09-13 | Safran Aircraft Engines | Module de chambre de combustion de turbomachine d'aeronef comprenant des marques facilitant le reperage lors d'une inspection endoscopique de la chambre de combustion |
CN112879163A (zh) * | 2021-01-11 | 2021-06-01 | 哈电发电设备国家工程研究中心有限公司 | 一种用于气路转换的新型气流分配转换装置 |
Citations (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0564171A1 (fr) * | 1992-03-30 | 1993-10-06 | General Electric Company | Capot en une seule pièce pour chambre annulaire double de combustion |
FR2704305A1 (fr) * | 1993-04-21 | 1994-10-28 | Snecma | Chambre de combustion comportant un système d'injection à géométrie variable. |
US5490378A (en) * | 1991-03-30 | 1996-02-13 | Mtu Motoren- Und Turbinen-Union Muenchen Gmbh | Gas turbine combustor |
EP0735318A2 (fr) * | 1995-03-25 | 1996-10-02 | Rolls-Royce Plc | Injecteur de carburant à géométrie variable |
US5934066A (en) * | 1996-10-18 | 1999-08-10 | Bmw Rolls-Royce Gmbh | Combustion chamber of a gas turbine with a ring-shaped head section |
WO1999063275A1 (fr) * | 1998-06-04 | 1999-12-09 | Pratt & Whitney Canada Corp. | Bande de refroidissement par gaine d'air pour chambre de combustion de moteur a turbine a gaz |
FR2909163A1 (fr) * | 2006-11-28 | 2008-05-30 | Snecma Sa | Carenage de chambre de combustion de turbomachine. |
FR2958015A1 (fr) * | 2010-03-24 | 2011-09-30 | Snecma | Systeme d'injection pour chambre de combustion de turbomachine, comprenant des moyens d'injection de carburant entre deux flux d'air coaxiaux |
Family Cites Families (10)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB947808A (en) * | 1961-08-03 | 1964-01-29 | Rolls Royce | Combustion chamber for a gas turbine engine |
US5323605A (en) * | 1990-10-01 | 1994-06-28 | General Electric Company | Double dome arched combustor |
CA2089272C (fr) * | 1992-03-23 | 2002-09-03 | James Norman Reinhold, Jr. | Chambre de combustion presentant une resistance aux chocs |
DE4313648C2 (de) * | 1993-04-21 | 1997-10-09 | Mannesmann Ag | Verfahren und Vorrichtung zum Herstellen von nahtlosen Rohren durch Drückwalzen |
FR2751054B1 (fr) * | 1996-07-11 | 1998-09-18 | Snecma | Chambre de combustion anti-nox a injection de carburant de type annulaire |
FR2856467B1 (fr) * | 2003-06-18 | 2005-09-02 | Snecma Moteurs | Chambre de combustion annulaire de turbomachine |
RU2250415C1 (ru) * | 2003-08-05 | 2005-04-20 | Открытое акционерное общество "Научно-производственное объединение "Сатурн" (ОАО "НПО "Сатурн") | Кольцевая камера сгорания газотурбинного двигателя |
US7975487B2 (en) * | 2006-09-21 | 2011-07-12 | Solar Turbines Inc. | Combustor assembly for gas turbine engine |
FR2943403B1 (fr) * | 2009-03-17 | 2014-11-14 | Snecma | Chambre de combustion de turbomachine comprenant des moyens ameliores d'alimentation en air |
FR2956897B1 (fr) * | 2010-02-26 | 2012-07-20 | Snecma | Systeme d'injection pour chambre de combustion de turbomachine, comprenant des moyens d'injection d'air ameliorant le melange air-carburant |
-
2013
- 2013-10-01 FR FR1359503A patent/FR3011317B1/fr active Active
-
2014
- 2014-09-29 EP EP14787226.1A patent/EP3039342B1/fr active Active
- 2014-09-29 RU RU2016116529A patent/RU2660729C2/ru active
- 2014-09-29 BR BR112016006688-0A patent/BR112016006688B1/pt active IP Right Grant
- 2014-09-29 CA CA2925441A patent/CA2925441C/fr active Active
- 2014-09-29 US US15/023,954 patent/US10180256B2/en active Active
- 2014-09-29 CN CN201480054491.6A patent/CN105593602B/zh active Active
- 2014-09-29 WO PCT/FR2014/052446 patent/WO2015049446A1/fr active Application Filing
Patent Citations (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5490378A (en) * | 1991-03-30 | 1996-02-13 | Mtu Motoren- Und Turbinen-Union Muenchen Gmbh | Gas turbine combustor |
EP0564171A1 (fr) * | 1992-03-30 | 1993-10-06 | General Electric Company | Capot en une seule pièce pour chambre annulaire double de combustion |
FR2704305A1 (fr) * | 1993-04-21 | 1994-10-28 | Snecma | Chambre de combustion comportant un système d'injection à géométrie variable. |
EP0735318A2 (fr) * | 1995-03-25 | 1996-10-02 | Rolls-Royce Plc | Injecteur de carburant à géométrie variable |
US5934066A (en) * | 1996-10-18 | 1999-08-10 | Bmw Rolls-Royce Gmbh | Combustion chamber of a gas turbine with a ring-shaped head section |
WO1999063275A1 (fr) * | 1998-06-04 | 1999-12-09 | Pratt & Whitney Canada Corp. | Bande de refroidissement par gaine d'air pour chambre de combustion de moteur a turbine a gaz |
FR2909163A1 (fr) * | 2006-11-28 | 2008-05-30 | Snecma Sa | Carenage de chambre de combustion de turbomachine. |
FR2958015A1 (fr) * | 2010-03-24 | 2011-09-30 | Snecma | Systeme d'injection pour chambre de combustion de turbomachine, comprenant des moyens d'injection de carburant entre deux flux d'air coaxiaux |
Also Published As
Publication number | Publication date |
---|---|
EP3039342B1 (fr) | 2018-04-04 |
FR3011317A1 (fr) | 2015-04-03 |
CN105593602B (zh) | 2018-07-24 |
FR3011317B1 (fr) | 2018-02-23 |
BR112016006688B1 (pt) | 2021-10-26 |
US20160215983A1 (en) | 2016-07-28 |
EP3039342A1 (fr) | 2016-07-06 |
CA2925441C (fr) | 2021-06-29 |
CN105593602A (zh) | 2016-05-18 |
CA2925441A1 (fr) | 2015-04-09 |
RU2660729C2 (ru) | 2018-07-09 |
BR112016006688A2 (pt) | 2017-08-01 |
RU2016116529A3 (fr) | 2018-05-08 |
US10180256B2 (en) | 2019-01-15 |
RU2016116529A (ru) | 2017-11-10 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
CA2754419C (fr) | Chambre de combustion de turbomachine comprenant des moyens ameliores d'alimentation en air | |
FR2930591A1 (fr) | Optimisation du positionnement angulaire d'un distributeur de turbine en sortie d'une chambre de combustion de turbomachine | |
EP3039341B1 (fr) | Chambre de combustion de turbomachine pourvue de moyens de déflection d'air pour réduire le sillage creé par une bougie d'allumage | |
EP3039342B1 (fr) | Chambre de combustion pour turbomachine a admission d'air homogène au travers de systèmes d'injection de carburant | |
WO2013060985A1 (fr) | Module de chambre de combustion de turbomachine d'aéronef et procédé de conception de celui-ci | |
WO2011054880A2 (fr) | Dispositif melangeur de carburant pour chambre de combustion de turbomachine comprenant des moyens ameliores d'alimentation en air | |
EP3530908B1 (fr) | Chambre de combustion comportant deux types d'injecteurs dans lesquels les organes d'étanchéité ont un seuil d'ouverture différent | |
FR2975465A1 (fr) | Paroi pour chambre de combustion de turbomachine comprenant un agencement optimise d'orifices d'entree d'air | |
FR2975467A1 (fr) | Systeme d'injection de carburant pour une chambre de combustion de turbomachine | |
CA2769342A1 (fr) | Chambre de combustion de turbomachine comprenant des orifices d'entree d'air ameliores | |
FR2952703A1 (fr) | Guide d'une bougie d'allumage dans une chambre de combustion d'une turbomachine | |
FR2958015A1 (fr) | Systeme d'injection pour chambre de combustion de turbomachine, comprenant des moyens d'injection de carburant entre deux flux d'air coaxiaux | |
FR2958373A1 (fr) | Chambre de combustion dans une turbomachine | |
EP3887720A1 (fr) | Systeme d'injection pour turbomachine, comprenant une vrille et des trous tourbillonnaires de bol melangeur | |
EP3449185B1 (fr) | Système d'injection de turbomachine comprenant un déflecteur aérodynamique à son entrée et une vrille d'admission d'air | |
FR3009747A1 (fr) | Chambre de combustion de turbomachine pourvue d'un passage d'entree d'air ameliore en aval d'un orifice de passage de bougie | |
FR2973479A1 (fr) | Paroi pour chambre de combustion de turbomachine comprenant un agencement optimise d'orifices d'entree d'air et de passage de bougie d'allumage | |
CA2850259C (fr) | Chambre de combustion annulaire dans une turbomachine | |
FR2948987A1 (fr) | Chambre de combustion de turbomachine comportant des orifices d'entree d'air ameliores | |
EP4004443A1 (fr) | Chambre de combustion comportant des systèmes d'injection secondaires et procédé d'alimentation en carburant | |
FR2957659A1 (fr) | Systeme d'injection pour chambre de combustion de turbomachine, comprenant des moyens d'injection de carburant en sortie d'une double vrille d'admission d'air | |
EP3969813B1 (fr) | Chambre de combustion comprenant des moyens de refroidissement d'une zone d'enveloppe annulaire en aval d'une cheminée | |
FR2952702A1 (fr) | Guidage d'une bougie d'allumage dans une chambre de combustion | |
FR2979005A1 (fr) | Systemes d'injection de carburant pour turbomachine d'aeronef a permeabilites differenciees | |
FR2943762A1 (fr) | Systeme d'injection de carburant dans une chambre de combustion de turbomachine |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
121 | Ep: the epo has been informed by wipo that ep was designated in this application |
Ref document number: 14787226 Country of ref document: EP Kind code of ref document: A1 |
|
WWE | Wipo information: entry into national phase |
Ref document number: 15023954 Country of ref document: US |
|
ENP | Entry into the national phase |
Ref document number: 2925441 Country of ref document: CA |
|
REEP | Request for entry into the european phase |
Ref document number: 2014787226 Country of ref document: EP |
|
WWE | Wipo information: entry into national phase |
Ref document number: 2014787226 Country of ref document: EP |
|
NENP | Non-entry into the national phase |
Ref country code: DE |
|
REG | Reference to national code |
Ref country code: BR Ref legal event code: B01A Ref document number: 112016006688 Country of ref document: BR |
|
ENP | Entry into the national phase |
Ref document number: 2016116529 Country of ref document: RU Kind code of ref document: A |
|
ENP | Entry into the national phase |
Ref document number: 112016006688 Country of ref document: BR Kind code of ref document: A2 Effective date: 20160328 |